Step-over blade-pitch control system

ABSTRACT

A pitch control system for blades on a rotor of an aircraft has a gimballing rotor hub ( 31 ) and a plurality of step-over arms ( 61 ) connected to the hub and capable of pivoting relative to the hub about a pivot axis. Each of a plurality of pitch links ( 55 ) connects one of the step-over arms ( 61 ) to a flight control system for pivoting the connected step-over arm ( 61 ) about the pivot axis and relative to the hub in response to inputs from the control system. Each of a plurality of step-over links ( 69 ) connects one of the step-over arms ( 61 ) to one of the blades for rotating the associated blade about the pitch axis in response to pivoting of the associated step-over arm.

TECHNICAL FIELD

The present application is related to blade-pitch control systems.

DESCRIPTION OF THE PRIOR ART

Tiltrotor aircraft have rotors that are moveable between a generallyvertical orientation for rotor-borne flight (helicopter mode) and agenerally horizontal orientation for wing-borne flight (airplane mode).One example of a tiltrotor aircraft is the Bell/Boeing V-22, which has apair of three-bladed rotors. To allow for use of a larger fuselage, morethrust, and/or higher speed, tiltrotors having four-bladed rotors havebeen proposed. However, four-bladed versions using prior-art types ofgimbaled rotor hubs can be unstable in airplane mode due to inadequatedamping of whirling.

Rotor-blade control systems for helicopters and tiltrotor aircraft arecomplex electrical and/or mechanical systems. The control systemsrespond to the pilot's input, but also must accommodate forces that actupon rotor assemblies and are generally outside the control of thepilot. Mechanical control systems typically include a swashplate, whichconsists of a stationary portion and a rotating portion. Typically, thelower, stationary portion is fixed in position and will not rotate, buthas the ability to move up and down and/or tilt in any given direction.This is commonly referred to as the “stationary” or “non-rotating”plate. Pilot inputs alter the vertical position of the stationary platethrough the collective control and the tilt of the stationary platethrough the cyclic control. The rotating portion of the swashplatearrangement is free to rotate. Pilot inputs to the non-rotating portionare passed through to the rotating portion of the control systems.

In the prior art, the rotating portion is typically connectedmechanically to each individual rotor blade. For example, in one type ofcontrol system, pitch links directly connect pitch horns on the rotorblades to the rotating plate of the swashplate, allowing the swashplateto alter the blade angle of each rotor blade.

However, it is necessary to include in control systems a subsystem whichreduces the degree of flapping as much as possible. In tiltrotoraircraft, it is especially important to counteract the detrimentaleffects of flapping, especially because the aircraft is capable of veryhigh speed travel, particularly in the airplane mode of flight. In theprior art, there are two basic approaches: one is to utilize an angledflap hinge; the other is to utilize offset pitch horns. Both of theseapproaches have the effect of introducing a kinematic pitch-flapcoupling, or delta-3, parameter in the system, and the delta-3 parameterrelates the amount of blade pitch change occurring for a given amount ofblade flapping motion. Designers seek to optimize delta-3 for counteringthe flapping encountered in flight.

Another kinematic coupling parameter which affects aeroelastic stabilityand rotor response of tiltrotors is the pitch-cone coupling, or delta-0,parameter. Like pitch-flap coupling, the pitch-cone coupling parameterrelates the amount of blade pitch change occurring for a given amount ofblade coning motion, which involves vertical motions of pairs of blades.The pitch-cone coupling caused by delta-0 alters the aerodynamic coningforces acting on the rotor which modifies the rotor response, rotorfrequency, and rotor hub forces. The pitch-cone coupling also changesthe sensitivity of the rotor system to gust disturbances and, in atiltrotor with four or more blades, can affect the flap-lag stability ofthe rotor system. This is because a tiltrotor with four or more bladeshas a reactionless coning mode, in which pairs of blades cone indifferent amounts and/or direction, that is not present on athree-bladed tiltrotor. The pitch-cone coupling alters the frequency ofthe out-of-plane reactionless coning mode frequency and can cause thismode to move closer to a reactionless in-plane mode. If the reactionlessconing mode frequency is too close to the reactionless in-plane modefrequency, then potential flap-lag instability may occur.

An optimized rotor hub design must provide the proper pitch-flapcoupling for controlling flapping and provide the proper pitch-conecoupling to ensure that flap-lag stability is maintained. Unfortunately,prior-art rotor hub configurations do not simultaneously provide desiredpitch-flap coupling and pitch-cone coupling and are compromiseconfigurations that optimize only one of the couplings.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is an oblique view of a tiltrotor aircraft having four-bladerotors.

FIG. 2 is an oblique view of a rotor hub assembly according to apreferred embodiment and used on the aircraft of FIG. 1.

FIG. 3 is a front view of the rotor hub of FIG. 2, the blade-pitchcontrol system of the assembly being shown in a reduced blade-pitchposition.

FIG. 4 is a side view of the rotor hub of FIG. 2, the blade-pitchcontrol system of the assembly being shown in a reduced blade-pitchposition.

FIG. 5 is a front view of the rotor hub of FIG. 2, the blade-pitchcontrol system of the assembly being shown in an increased blade-pitchposition.

FIG. 6 is a side view of the rotor hub of FIG. 2, the blade-pitchcontrol system of the assembly being shown in an increased blade-pitchposition.

FIG. 7 is a front view of the rotor hub of FIG. 2, the blade-pitchcontrol system of the assembly being shown in a reduced blade-pitchposition, the hub assembly being shown in a gimbaled orientation.

FIG. 8 is a graph showing the relationship between blade pitch and thedelta-3 angle for two values of pre-cone for the rotor hub of FIG. 2.

DETAILED DESCRIPTION

A gimbaled rotor hub configuration is provided for use on an aircraft,the rotor hub configuration being particularly useful on tiltrotoraircraft. The rotor hub has a gimbaled yoke, which allows for flappingmotions, and the blades of the rotor are adjustable for pitch angle. Astep-over linkage between the swashplate and the pitch horn provides forcontrol of the blade pitch, and this step-over linkage is able tosimultaneously provide a desired value of pitch-flap coupling (delta-3)for whirl flutter stability and a desired value of pitch-cone coupling(delta-0) for reaction-less flap-lag stability. The step-over linkageallows a rotor having four or more blades to have the same delta-3 valueas a three-blade rotor, can provide for delta-3 that varies withcollective input to increase stability, and allows for the selection ofdesired delta-0 values. Without the step-over mechanism, one of thesecoupling parameters will not be at an optimum setting and will result inreduced aeroelastic stability.

Referring to FIG. 1, aircraft 11 is a tiltrotor aircraft having afuselage 13 and wings 15, 17 extending from fuselage 13. A nacelle 19,21 is rotatably mounted to the outer end of each wing 15, 17, and eachnacelle 19, 21 houses an engine (not shown) for rotating an attachedprop-rotor 23, 25. Each prop-rotor, or rotor, has a plurality of rotorblades 27, and the embodiment shown has four blades 27 per rotor 23, 25.Each rotor 23, 25 also has a central rotor hub, which retains blades 27and is located under a spinning cover 29. The rotor hubs are gimbaledhubs and have a step-over linkage, which is described below.

FIGS. 2 through 7 show a preferred embodiment of a rotor hub used on theaircraft of FIG. 1. Rotor hub assembly 31 comprises a yoke 33, which isconnected to a mast 35 by a constant-velocity torque coupling 37 forrotation with the mast. In the embodiment shown yoke comprises fourblade attachment arms 39 a, 39 b, 39 c, 39 d, and yoke is rigidlyconnected to torque coupling 37 with fasteners. Torque coupling 37 has aportion that is pivotable relative to mast 35 through rotation aboutperpendicular axes 43, 45 on bearings 47, and this configuration allowsfor yoke and blades (not shown) attached to arms 39 a, 39 b, 39 c, 39 dto gimbal relative to mast 35. As an illustrative example, a blade grip49 is shown attached to arm 39 d, blade grip 49 being pivotable on arm39 d about pitch axis for adjustability of the pitch angle of anattached blade. Blade grip 49 has a pitch horn 53 located on an inboardend of grip 49 and extending radially from grip 49. Movement of pitchhorn 53 about axis 51 causes a corresponding change in blade pitchangle. Hub assembly 31 is shown with only one blade grip 49 on yoke 33,though a blade grip 49 and blade would be attached to each of arms 39 a,39 b, 39 c, 39 d in the complete assembly.

Step-over linkages (only one shown) are provided for connecting eachpitch horn 53 to a flight control system (not shown), such as, forexample, a swashplate, for controlling the pitch angle of blades inresponse to input from the flight control system. The flight controlsystem may be in a fixed position relative to mast 35 or may moverelative to mast during operation, but the flight control system doesnot gimbal with yoke 33 and the attached blades relative to mast. Arigid pitch link 55 has spherical bearing rod ends 57, 59 on oppositeends of link 55, with a lower rod end 57 being connected to the flightcontrol system, and an upper rod end 59 being connected to a step-overarm 61. Step-over arm 61 is a rigid member and is pivotally connected ata root end 63 to a step-over mount 65, which is rigidly connected totorque coupling 37. Each step-over arm pivots about a step-over axis 66.A link end 67, which is opposite root end 63, is configured forreceiving rod end 59 of pitch link 55. A rigid step-over link 69connects link end 67 of step-over arm 61 with pitch horn 53 of bladegrip 49, step-over link 69 having lower rod end 71 and upper rod end 73on opposite ends of link 69, each rod end 71, 73 being a spherical rodend. Use of spherical-bearing rod ends 57, 59, 71, 73 allows links 55,69 to pivot relative to the component connected at each end of links 55,69 at varying angles without interference.

The step-over linkage provides a significant advantage due to thedecoupling of pitch-flap (delta-3) and pitch-cone (delta-0) kinematicparameters. This is due to the fact that the delta-0 term derives fromthe angle formed between a pitch horn of each blade, such as pitch horn53, and the coning axis of the blades, whereas the delta-3 term derivesfrom the angle formed between pitch link 55 and flapping axes 43, 45. Afour-bar linkage is formed from pitch horn 53/blade grip 49, step-overlink 69, step-over arm 61, and torque coupling 37/yoke 33, and thisfour-bar linkage gimbals relative to mast 35 and relative to pitch link55 during flapping. This means that the angles between members of thefour-bar linkage do not change due to flapping, and the only anglechange caused by flapping is between pitch link 55 and link end 67 ofstep-over arm 61. Another advantage to using the step-over linkage isthat pitch link 55 can be located further toward an adjacent rotorblade, to achieve desirable coupling terms, than would be possiblewithout interference when using one link extending between the flightcontrol system and pitch horn 53.

FIGS. 3 through 6 show rotor hub assembly 31 with the step-over linkagemoved between a first position (FIGS. 3 and 4), which corresponds toreduced blade pitch, and a second position (FIGS. 5 and 6), whichcorresponds to increased blade pitch. In the preferred embodiment, thestep-over linkage moves in a direction opposite that of the leading edgeof the associated blade due to pitch horn 53 being located in a trailingposition on blade grip 49.

FIGS. 3 and 4 show hub assembly 31 with pitch link 55 moved to anuppermost position, causing step-over arm 61 to rotate about axis 66 andstep-over link 69 to also be moved into an uppermost position. Pitchhorn 53 is thus moved into an uppermost position, causing blade grip 49to rotate about pitch axis 51, such that the attached blade is moved toa minimum blade-pitch position.

FIGS. 5 and 6 show hub assembly 31 with pitch link 55 moved to alowermost position, causing step-over arm 61 to rotate about axis 66 andstep-over link 69 to also be moved into a lowermost position. Pitch horn53 is thus moved into a lowermost position, causing blade grip 49 torotate about pitch axis 51, such that the attached blade is moved to amaximum blade-pitch position.

FIG. 7 shows hub assembly 31 with step-over linkage moved into theuppermost position and with hub assembly 31 gimbaled relative to mast35. Assembly 31 is shown rotated about an axis defined by bearing 47, aswould occur during flapping, in which advancing blades rise andretreating blades fall. Flapping is allowed to gimbal hub assembly 31,such that yoke 33 and attached blades (not shown) are at an anglerelative to mast 35. As described above, the figure shows that therelative angles of the step-over linkage do not change during flapping,except for the angle between pitch link 55 and step-over arm 61. This iswhere, the unflapped system (pitch link 55 and the associated flightcontrol system) meets the flapped system (hub assembly 31) and where thedelta-3 parameter is determined. In this manner, the delta-0 parameteris decoupled from the delta-3 parameter.

The step-over linkage is particularly useful for tiltrotor aircraft. Ahigh delta-3 is desirable in helicopter mode, in which blades arepositioned at a small angle relative to the rotor plane, and a lowdelta-3 is desirable in airplane mode (low collective), in which bladesare positioned at a large angle relative to the rotor plane (highcollective). Prior-art blade-pitch linkage configurations required anundesirable compromise to be made for one or more parameters. Becausethe delta-0 and delta-3 parameters are decoupled in the step-overlinkage, coning does not affect delta-3, and delta-3 can be optimizedthroughout the range of collective. This is accomplished by angling thestep-over axis 66 relative to the rotor plane, as can be seen in FIGS.3, 5 and 7. Step-over axis 66 is shown as being angled slightlydownward.

FIG. 8 is a graph showing the plots of negative delta-3 angle versus agiven blade pitch angle for two values of precone, which is the angle ofstep-over axis 66. As described above, a high negative delta-3 isdesirable at low collective, which corresponds to helicopter mode and islocated toward the left on the x-axis, and a low value of negativedelta-3 is desirable at high collective, which corresponds to airplanemode and is located toward the right side on the x-axis. As shown in thegraph, for a range of blade pitch from −8 degrees to +54 degrees, aprecone value of −5 degrees provides for a range of approximately −31degrees to approximately −16 degrees of delta-3 angle. However, a lowervalue of negative delta-3 is typically desirable in airplane mode. Thevalues plotted for −10 degrees of precone show that negative delta-3 isslightly higher in helicopter mode and that delta-3 continues to improveas blade pitch increases, providing favorable coupling terms throughoutthe range of blade pitch.

It should be noted that the relative locations, as shown, of componentsin the step-over linkage are to be considered examples. The step-overlinkage may be altered from the configuration shown to provide forvarious advantageous qualities or parameters. For example, through pitchlink 55 and step-over link 69 are shown as being connected to step-overarm 61 at approximately the same location, links 55, 69 may be connectedat different distances from step-over axis 66. This would allow for areduction or increase in the amount of travel of one link 55, 69relative to the other link 55, 69.

The step-over linkage configuration provides for several advantages,including: (1) providing a simple control system for controlling pitchof blades on a gimbaled rotor; (2) providing decoupled pitch-flap andpitch-cone kinematics; and (3) providing the capability for desirablevalues for both pitch-flap and pitch-cone coupling.

Though reference is made to an illustrative embodiment, this descriptionis not intended to be construed in a limiting sense. Variousmodifications and combinations of the illustrative embodiments, as wellas other embodiments, will be apparent to persons skilled in the artupon reference to the description.

1. A blade-pitch control system for an aircraft, the control systemcontrolling motion of each of a plurality of rotor blades about anassociated pitch axis, the control system comprising: a rotor hubadapted for rotation with a mast and for gimballing relative to themast; a plurality of step-over arms, each step-over arm being connectedto the hub and capable of pivoting relative to the hub about a pivotaxis that is in a fixed orientation relative to the hub and moves withthe hub during gimballing of the hub; a plurality of pitch links, eachpitch link being adapted for connecting one of the step-over arms to aflight control system for pivoting the connected step-over arm relativeto the hub in response to inputs from the flight control system; and aplurality of step-over links, each step-over link being adapted forconnecting one of the step-over arms to one of the blades for rotatingthe associated blade about the corresponding pitch axis in response topivoting of the associated step-over arm; wherein each step-over arm isan elongated arm, one end of the step-over arm being pivotally connectedto the hub and an opposite end being free to rotate about the pivotaxis, the associated pitch link and step-over link being connected tothe step-over arm at a location toward the free end of the step-overarm.
 2. The control system according to claim 1, wherein each pivot axisis oriented at a selected angle relative to the hub.
 3. The controlsystem according to claim 1, wherein the flight control system comprisesa swashplate, and the pitch links are adapted to connect the step-overarms to a rotatable portion of the swashplate.
 4. A rotor hub assembly,comprising: a coupling adapted for rotation with a mast, at least aportion of the coupling being configured for gimballing relative to themast; a yoke connected to the coupling and adapted to carry a pluralityof blades, each blade being pivotable about a pitch axis, the yoke beingconfigured to move with the coupling during gimballing of the coupling;a plurality of step-over arms, each step-over arm being connected to thecoupling and capable of pivoting relative to the coupling about a pivotaxis that is in a fixed orientation relative to the coupling and moveswith the coupling during gimballing of the coupling; a plurality ofpitch links, each pitch link being adapted for connecting one of thestep-over arms to a flight control system for pivoting the connectedstep-over arm relative to the coupling in response to inputs from theflight control system; and a plurality of step-over links, eachstep-over link being adapted for connecting one of the step-over arms toone of the blades for rotating the associated blade about thecorresponding pitch axis in response to pivoting of the associatedstep-over arm; wherein each step-over arm is an elongated arm, one endof the step-over arm being pivotally connected to the coupling and anopposite end being free to rotate about the pivot axis, the associatedpitch link and step-over link being connected to the step-over arm at alocation toward the free end of the step-over arm.
 5. The control systemaccording to claim 4, wherein each pivot axis is oriented at a selectedangle relative to the coupling.
 6. The control system according to claim4 wherein the flight control system comprises a swashplate, and thepitch links are adapted to connect the step-over arms to a rotatableportion of the swashplate.
 7. A rotary-wing aircraft, comprising: amast; a coupling configured for rotation with the mast, at least aportion of the coupling being configured for gimballing relative to themast; a yoke connected to the coupling; the yoke being configured togimbal with the coupling; a plurality of blades carried by the yoke,each blade being pivotable about a pitch axis: a plurality of step-overarms, each step-over arm being connected to the coupling and capable ofpivoting relative to the coupling about a pivot axis that is in a fixedorientation relative to the coupling and gimbals with the coupling; aplurality of pitch links, each pitch link being adapted for connectingone of the step-over arms to a flight control system for pivoting theconnected step-over arm relative to the coupling in response to inputsfrom the flight control system; and a plurality of step-over links, eachstep-over link being adapted for connecting one of the step-over arms toone of the blades for rotating the associated blade about thecorresponding pitch axis in response to pivoting of the associatedstep-over arm; wherein each step-over arm is an elongated arm, one endof the step-over arm being pivotally connected to the coupling and anopposite end being free to rotate about the pivot axis, the associatedpitch link and step-over link being connected to the step-over arm at alocation toward the free end of the step-over arm.
 8. The control systemaccording to claim 7, wherein each pivot axis is oriented at a selectedangle relative to the coupling.
 9. The control system according to claim7, wherein the flight control system comprises a swashplate, and thepitch links are adapted to connect the step-over arms to a rotatableportion of the swashplate.